1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a large air cooled turbine blade.
2. Description of the Related Art Including, Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine of the type used in electrical power production includes a turbine section with three or four stages or rows of rotor blades. The last stage rotor blades are very large. The prior art cooling of a large turbine rotor blade is achieved by drilling radial holes into the blade from the tip and root sections. Limitation of drilling a long radial hole from both ends of the airfoil increases for a large and highly twisted and tapered blade airfoil. Reduction of the available airfoil cross section area for drilling radial holes is a function of the blade twist and taper. Higher airfoil twist and taper yield a lower available cross sectional area for drilling radial cooling holes. Cooling of the large and highly twisted and tapered blade by this manufacturing process will not achieve the optimum blade cooling effectiveness. Especially effective cooling for the airfoil leading and trailing edges are difficult to achieve. the prior art process for producing large and highly twisted turbine blades prevent a blade that can be used in a high temperature environment or with the use of low cooling flow, both of which the future requires for next generation industrial gas turbine engines.